With reference to FIG. 1, a ducted fan gas turbine engine generally indicated at 10 has a principal and rotational axis X-X. The engine comprises, in axial flow series, to an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, and intermediate-pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 14 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
The performance of gas turbine engines, whether measured in terms of efficiency or specific output, is improved by increasing the turbine gas temperature. It is therefore desirable to operate the turbines at the highest possible temperatures. For any engine cycle compression ratio or bypass ratio, increasing the turbine entry gas temperature produces more specific thrust (e.g. engine thrust per unit of air mass flow). However as turbine entry temperatures increase, the life of an un-cooled turbine falls, necessitating the development of better materials and the introduction of internal air cooling.
In modern engines, the high-pressure turbine gas temperatures are hotter than the melting point of the material of the blades and vanes, necessitating internal air cooling of these airfoil components. During its passage through the engine, the mean temperature of the gas stream decreases as power is extracted. Therefore, the need to cool the static and rotary parts of the engine structure decreases as the gas moves from the high-pressure stage(s), through the intermediate-pressure and low-pressure stages, and towards the exit nozzle.
FIG. 2 shows an isometric view of a typical single stage cooled turbine. Cooling air flows are indicated by arrows.
Internal convection and external films are the prime methods of cooling the gas path components—airfoils, platforms, shrouds and shroud segments etc. High-pressure turbine nozzle guide vanes 31 (NGVs) consume the greatest amount of cooling air on high temperature engines. High-pressure blades 32 typically use about half of the NGV flow. The intermediate-pressure and low-pressure stages downstream of the HP turbine use progressively less cooling air.
The high-pressure turbine airfoils are cooled by using high pressure air from the compressor that has by-passed the combustor and is therefore relatively cool compared to the gas temperature. Typical cooling air temperatures are between 800 and 1000 K, while gas temperatures can be in excess of 2100 K.
The cooling air from the compressor that is used to cool the hot turbine components is not used fully to extract work from the turbine. Therefore, as extracting coolant flow has an adverse effect on the engine operating efficiency, it is important to use the cooling air effectively.
Ever increasing gas temperature levels combined with a drive towards flatter combustion radial profiles, in the interests of reduced combustor emissions, have resulted in an increase in local gas temperature experienced by the working gas annulus endwalls, which include NGV platforms 33, blade platforms 34 and shroud segments 35 (also known as shroud liners). However, the flow of air that is used to cool these endwalls can be highly detrimental to the turbine efficiency. This is due to the high mixing losses attributed to these cooling flows when they are returned to the mainstream working gas path flow.
One option is to cool the platforms and shroud segments by an impingement flow of cooling air on the back plane of the gas washed wall of the component. For example, a perforated plate spaced from the gas washed wall and supported by pedestals can form impinging jets, and the spent coolant can then flow back into the working gas path at the rear edges of the component. Unfortunately, limited numbers of impingement jets can produce non uniform heat transfer distributions, and the cross flow from spent coolant can reduce the effectiveness of the impingement jets at the more downstream locations of the component. In addition, the need to keep the coolant pressure at a level above that in the working gas path reduces the allowable pressure drop across the impingement jets, and hence the associated heat transfer levels. Further, most impingement schemes involve brazing or laser welding a sheet metal plate onto a cast component. The drilling of impingement holes into this plate adds cost, and the plate can become detached if the braze or weld cracks in the field due to thermal fatigue
Thus, another option is to provide one or more internal cooling passages behind the endwall. Such passages are typically bounded between spaced upper and lower walls that follow the shape of the endwall. As shown in FIG. 3(a), one arrangement has multi-pass passages 40, optionally augmented with heat transfer augmentation features such as trip strips 41. Another arrangement has a wider passage 42 augmented with e.g. an array of staggered pedestals 43 (FIG. 3(b) or parallel fins 44 (FIG. 3(c) to increase the wetted plane area of the passage. In FIGS. 3(a)-(c) the block arrows indicate cooling air flows. Conveniently, such arrangements can usually be formed using single pull die ceramic core technology. However, the arrangements suffer from various limitations. For example, the multi-pass passage arrangement provides a relatively low wetted plane area, the wider passage with a staggered pedestal array also has a relatively low wetted plane area, and the wider passage with parallel fins can suffer from a low level of heat transfer as a boundary layer thickens along the channels formed by the fins. Coolant flow levels therefore have to be increased to offset the low wetted plane area and/or low level of heat transfer. Further, the heat transfer augmentation features can provide unreliable flow distributions, which in turn can lead to unreliable heat transfer coefficients.